A stationary blade of a gas turbine used for generating electric power provides a flow passage for combustion gas, of which temperature reaches about 1300° C. Therefore, in order to prevent melt or damage of the stationary blade by combustion gas, various cooling structures are provided to the gas turbine stationary blade. As a technique concerning such a cooling structure, there is a widely-known technique in which a passage for cooling air is provided in the gas turbine stationary blade, cooling air is sent into this passage, thereby cooling the gas turbine stationary blade from inside (refer to Japanese Patent Application Laid-open No. 11-132005 for example).
The gas turbine stationary blade has a dividable structure capable of assembling and disassembling the gas turbine stationary blade in consideration of easy maintenance after installation thereof. FIG. 14 is a perspective view showing a segment 1 that is a unit constituent element of a two-staged stationary blade of a gas turbine. Each unit constituting this segment 1 comprises a substantially parallelogram inner shroud 2, one columnar stationary blade section 3 whose one end is fixed to the inner shroud 2, and a substantially parallelogram outer shroud 4 installed substantially in parallel to the inner shroud 2 and fixed to the other end of the stationary blade section 3. The segment 1 comprises a pair of the units welded and connected to each other side-by-side. The gas turbine stationary blade comprises a plurality of segments 1 connected to one other side-by-side through detachable connection members (not illustrated) such as bolts such that the gas turbine stationary blade is formed into an annular structure as a whole. The gas turbine stationary blade is fixed and installed in a gas turbine casing (not illustrated) with a cantilever structure by means of legs 5 provided on an outer peripheral side face of the outer shroud 4.
A bolt joint section 7 of the segment 1 keeps a specific distance so as to absorb expansion of the gas turbine stationary blade when the gas turbine is driven. This distance is set such that the distance is made zero by expansion of the gas turbine stationary blade when the gas turbine is driven. However, due to tolerance during the actual producing procedure, a gap 7a ranging from about 0.5 mm to 1 mm is produced in the bolt joint section 7.
FIG. 15 is an enlarged perspective view around the inner shroud 2 shown in FIG. 14. FIG. 16 is a plan sectional view of the inner shroud 2 shown in FIG. 15. FIG. 17 is a side sectional view of the inner shroud 2 taken along the line I—I in FIG. 16. FIG. 18 is a side sectional view of the inner shroud 2 taken along the line II—II in FIG. 16. In FIG. 15 to FIG. 18, the gas turbine stationary blade has a stationary blade section front edge passage 9 and a stationary blade section rear edge passage 10 isolated from each other by a rib 8 which are provided inside the stationary blade section 3. The stationary blade section front edge passage 9 is in communication with an open chamber 11 provided in the inner shroud 2. The stationary blade section rear edge passage 10 passes through the inner shroud 2, and is in communication with a cavity 12 formed in a bottom face section of the inner shroud 2. The open chamber 11 and the cavity 12 are isolated from each other by a bottom plate 13 installed on the bottom face section of the inner shroud 2. A member 14, shown in FIG. 17, in the stationary blade section 3 is an impingement tube 14 comprising a metal member inserted into the stationary blade section front edge passage 9 and the stationary blade section rear edge passage 10 so as to subject the stationary blade section 3 to impingement cooling.
In the inner shroud 2, a front edge 15 is located in upstream portion in the flow passage for combustion gas 6. A front edge flow passage 16 is provided along the front edge 15. The front edge flow passage 16 and the open chamber 11 are in communication with each other through an intermediate flow passage 17 provided therebetween. A regulating plate 18 is laid on a floor section of the front edge flow passage 16 to narrow a cross sectional area of the flow passage. A plurality of turbulators 20 are provided on the regulating plate 18 and a ceiling section of the front edge flow passage 16 to agitate the cooling air 19.
From an outlet orifice of the front edge flow passage 16, a central flow passage 21 having a cross sectional area smaller than that of the front edge flow passage 16 is pulled out. The central flow passage 21 comes out from a rear edge 23 of the inner shroud 2 that is downstream of the flow passage of combustion gas along the welded joints 22 of the inner shrouds 2. From a position near an inlet orifice of the front edge flow passage 16 also, a side edge flow passage 24 having a cross sectional area smaller than that of the front edge flow passage 16 is pulled out. The side edge flow passage 24 comes out from the rear edge 23 along a side edge 25 of the inner shroud 2 (refer to FIG. 16 and FIG. 18). The cooling structure is provided for each pair of units constituting the segment 1, and a pair of left and right cooling structures are provided to constitute a cooling structure of the inner shroud 2.
At the time of actuation of the gas turbine, when the inner shroud 2 is to be cooled, cooling air 19 is sent to the impingement tube 14 in the stationary blade section 3 from the outer shroud 4. The cooling air 19 subjects the stationary blade section 3 to impingement cooling, a portion of the cooling air 19 flows into the open chamber 11 in the inner shroud 2 through the front edge passage 9 of the stationary blade section, and a portion of the cooling air 19 penetrates the inner shroud 2 through the stationary blade section rear edge passage 10 and is supplied to the cavity 12 (refer to FIG. 17). The cooling air 19 which has flowed into the open chamber 11 flows into the front edge flow passage 16 through the intermediate flow passage 17 to convection-cools the front edge of the inner shroud 2. A portion of the air flows into the side edge flow passage 24 from an inlet orifice of the front edge flow passage 16, convection-cools the side edge 25 of the inner shroud 2, and is discharged from the rear edge 23. Remaining cooling air 19 flows into the central flow passage 21 from the outlet orifice of the front edge flow passage 16, convection-cools welded joints 22 of the inner shrouds, and is discharged from the rear edge 23.
The regulating plate 18 is provided to prevent reduction in flow speed of the cooling air 19 which passes through the front edge flow passage 16 by narrowing the cross sectional area of the passage, and to enhance the cooling efficiency of the front edge 15. The turbulator 20 agitates the cooling air 19 in the front edge flow passage 16, and enhances cooling efficiency of the front edge 15. The central flow passage 21 and the side edge flow passage 24 have cross sectional areas smaller than those of the front edge flow passage 16. Therefore, flow speed of the cooling air 19 passing through the flow passages 21 and 24 is faster than that in the front edge flow passage 16. Thus, the structure in which the flow passage is narrowed enhances the cooling efficiency near the welded joints 22 of the inner shrouds 2 and near the side edge 25.
The cooling air 19 supplied to the cavity 12 is used as sealing air for sealing a gap (not illustrated) between the gas turbine stationary blade and a gas turbine rotor blade. A portion of the sealing air is blown out from a bottom surface section of the front edge 15, to film-cool the inner shroud 2 from the front edge 15.
According to research of the present inventors, however, there has been found a phenomenon that corner sections of the front edge 15 of the inner shroud 2 in the conventional gas turbine stationary blade are burned 26, 26 and that thermal barrier coating (TBC) which is applied to the surface is removed (refer to FIG. 15). It is conceived that this is because in the conventional technique, the flow passages 16 and 24 of the cooling air 19 are not in communication up to the corner section of the front edge 15 and that the corner section is insufficiently cooled. Further, in the conventional technique, it is conceived that since both the flow passages of the central flow passage 21 and the side edge flow passage 24 are pulled out from the edge flow passage 16, the flow speed of the cooling air 19 passing through the flow passages 21 and 24 is reduced to cause reduction of the cooling efficiency.
A method in which the regulating plate 18 and the turbulator 20 are provided to maintain the cooling efficiency like the conventional technique has a problem that the cooling efficiency of the corner section of the front edge 15 is inferior and the gas turbine structure is complicated to increase production costs. A method in which the flow rate of the cooling air 19 in the inner shroud 2 is largely increased to enhance the cooling efficiency has a problem that driving cost of the gas turbine increases. There is another method in which the cross sectional areas of the central flow passage 21 and the side edge flow passage 24 are narrowed to increase the flow speed of the cooling air 19, but if the cross sectional area of the flow passage is excessively narrowed, the range that can be cooled is also narrowed, and therefore the cooling efficiency of the corner section of the front edge 15 is further lowered.